Cooling of gas turbine engine accessories

ABSTRACT

A gas turbine engine for an aircraft is provided. The engine includes an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor. The engine further includes core casings surrounding the engine core. The engine further includes an aerodynamic cowl which surrounds the core casings. The engine further includes a propulsive fan located upstream of the engine core, the fan generating a core airflow which enters the core engine and a bypass airflow which enters a bypass duct surrounding the aerodynamic cowl. The engine further includes one or more engine accessories mounted in a space between the core casings and the aerodynamic cowl.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is based upon and claims the benefit of priority fromUnited Kingdom Patent Application No. GB 1817842.6, filed on 1 Nov.2018, which is hereby incorporated herein in its entirety.

BACKGROUND Technical Field

The present disclosure relates to cooling of engine accessories of a gasturbine engine.

Description of the Related Art

Conventionally the accessory gearbox of an aircraft gas turbine engineis mounted within the outer engine nacelle in a location beneath theengine. The gearbox is connected to the engine core by a radial driveshaft. The gearbox provides power to other accessories such as anauxiliary generator and pumps for hydraulic fluid, fuel, oil etc.

Although effective in that the gearbox is kept away from the hotenvironment of the engine core, the location of the gearbox within theouter nacelle is disadvantageous in that it requires a relativelysignificant amount of space within the nacelle, which can increase theoverall diameter of the nacelle, leading to weight and drag increase andadverse specific fuel consumption.

In alternative arrangements, it is possible to locate the accessorygearbox and other accessories in an engine zone directly outside thecasings which surround the engine core. The accessories are mounted tothe casings by brackets, short spars or bosses. However, the hightemperature environment near the engine core can produce thermal loadswhich can reduce component reliability. For example, in some engines thezone can operate at air temperatures above 300° C., which is above thethermal capability of many accessory gearbox materials. Therefore tomaintain acceptable surrounding air temperatures for the gearbox, aventilation system may be installed.

SUMMARY

The present disclosure is at least partly based on a realisation that apossible source of cooling air in turbofan engines is the bypass airflowthrough the bypass duct of the engine. This airflow is typically attemperatures below 100°. One option, therefore, is to form simpleventilation holes in the aerodynamic cowl which surrounds the enginezone so that a small portion of the bypass airflow enters the zone tosurround the accessory gearbox. However, in high bypass ratio engines,particularly at low engine power conditions, a low fan pressure ratioreduces the pressure difference across the ventilation holes and therebylimits the amount of the bypass flow that can be diverted in this way tocool the gearbox. This in turn limits the capability of the ventilationsystem.

Thus according to a first aspect there is provided a gas turbine enginefor an aircraft, the engine including:

an engine core comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor;

core casings surrounding the engine core;

an aerodynamic cowl which surrounds the core casings;

a propulsive fan located upstream of the engine core, the fan generatinga core airflow which enters the core engine and a bypass airflow whichenters a bypass duct surrounding the aerodynamic cowl; and

one or more engine accessories mounted in a space between the corecasings and the aerodynamic cowl;

wherein the gas turbine engine further includes:

one or more ventilation inlets which receive a portion of the bypassairflow as a cooling flow for the engine accessories, the ventilationinlets being configured to convert at least a portion of the kineticenergy of the bypass flow received therein into a pressure rise suchthat the pressure of the cooling flow is increased relative to thepressure of the bypass airflow; and

a manifold in fluid communication with the ventilation inlets to collectthe cooling flow therefrom, the manifold having plural exhaust holestherefrom through which the cooling flow leaves the manifold to impingeupon and thereby cool the engine accessories.

Advantageously, by increasing the pressure of the cooling flow in thisway, and directing it to where it is needed using the exhaust holes ofthe manifold, the problem of a low fan pressure ratio can be overcomesuch that even at low engine power conditions sufficient cooling of theengine accessories can be achieved.

Optional features of the present disclosure will now be set out. Theseare applicable singly or in any combination with any aspect of thepresent disclosure.

The one or more ventilation inlets may convert at least 50% of thekinetic energy of the portion of the bypass airflow received thereininto a pressure rise. Preferably they may substantially stagnate theportion of the bypass airflow received therein.

The gas turbine engine may further include a control unit which isoperable to regulate the amount of cooling flow collected by themanifold from the ventilation inlets. In this way, the amount of coolingflow can be adapted to different flight conditions, reducing losses andimproving engine efficiency. For example, the control unit can be underthe control of an engine electronic controller (EEC) which manages theoperation of the engine as a whole. In contrast, simple ventilationholes formed in the aerodynamic cowl would not allow the amount ofcooling flow to be adapted in this way so that, for example, holes sizedfor the maximum power take-off condition where the engine core operatingtemperatures are at their highest would tend to provide larger thannecessary ventilation flows at engine cruise, negatively impacting theperformance of the engine and increasing the engine specific fuelconsumption.

The exhaust holes may be arranged such that the cooling flow impingingupon the engine accessories forms a continuous film of cooling airaround the accessories.

Conveniently, the exhaust holes may be formed into two lines whichextend substantially parallel to the principal rotation axis of theengine to opposite sides of the engine accessories.

The one or more engine accessories may be mounted vertically beneath thecore casings.

The one or more engine accessories may include an engine accessorygearbox driven by a take-off (e.g. radial drive shaft) from the coreshaft. Other of the engine accessories can include, for example, any oneor more of an electrical power generator, a fuel pump, an oil pump, ahydraulic pump, and an engine starter motor.

The engine accessory gearbox may include a train of spur gears whichtransfer the drive to other engine accessories, the spur gears beingarranged in a line and having axes of rotation which extendperpendicularly to the principal rotation axis of the engine. Forexample, the spur gears of the train may be arranged in a line withtheir axes of rotation extending perpendicularly to the engine'sprincipal rotation axis. In particular, the line may extend in adirection which is substantially parallel with the engine axis. This isin contrast with many conventional engine accessory gearboxes, where thetrain of spur gears extend around a circumferential direction of theengine. The train of spur gears may be mounted along a central spinemember, the other engine accessories projecting from opposite sides ofthe spine member.

The gas turbine engine may further include a circumferential row ofouter guide vanes located in the bypass duct rearwards of the propulsivefan, the outer guide vanes extending radially outwardly from an innerring which defines a radially inner surface of the bypass duct. Theaerodynamic cowl can then be rearwards of the inner ring and can includetwo door sections located on respective and opposite sides of theengine, each door section being pivotably openable about an upper edgethereof to enable maintenance access to the engine core, and theaerodynamic cowl can further include a keel beam which extendsrearwardly from the inner ring at bottom dead centre thereof to providelatching formations for latching to a lower edge of each door sectionwhen it is closed. Conveniently, in such an arrangement, the ventilationinlets may be located on the radially outer surface of the keel beam.The manifold may be located on the radially inner surface of the keelbeam. The keel beam, as well as providing a convenient location for theventilation inlets and/or manifold, allows the lower edges of the doorsections to be located away from bottom dead centre where core engineliquids are most likely to accumulate. Instead these liquids can beguided and drained by appropriate shaping of the inner surface of thekeel beam. Also, the keel beam can provide a relative rigid structurefor sealing to the lower edges of the door sections, thereby improvingthese seals and hence the fire zone boundary performance between coreand bypass zones of the engine.

The gas turbine engine may further include a power gearbox that receivesan input from the core shaft and outputs drive to the propulsive fan soas to drive the fan at a lower rotational speed than the core shaft.

The turbine may be a first turbine, the compressor may be a firstcompressor, and the core shaft may be a first core shaft. The enginecore may then further include a second turbine, a second compressor, anda second core shaft connecting the second turbine to the secondcompressor. The second turbine, second compressor, and second core shaftmay be arranged to rotate at a higher rotational speed than the firstcore shaft.

As noted elsewhere herein, arrangements of the present disclosure may beparticularly, although not exclusively, beneficial for fans that aredriven via a gearbox. Accordingly, the gas turbine engine may comprise agearbox that receives an input from the core shaft and outputs drive tothe fan so as to drive the fan at a lower rotational speed than the coreshaft. The input to the gearbox may be directly from the core shaft, orindirectly from the core shaft, for example via a spur shaft and/orgear. The core shaft may rigidly connect the turbine and the compressor,such that the turbine and compressor rotate at the same speed (with thefan rotating at a lower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages.

Each stage may comprise a row of rotor blades and a row of stator vanes.The row of rotor blades and the row of stator vanes may be axiallyoffset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). These ratios may commonly be referred to as the hub-to-tipratio. The radius at the hub and the radius at the tip may both bemeasured at the leading edge (or axially forwardmost) part of the blade.The hub-to-tip ratio refers, of course, to the gas-washed portion of thefan blade, i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 250 cm (around 100 inches), 260 cm, 270 cm (around 105inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm(around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around130 inches), 340 cm (around 135 inches), 350 cm, 360 cm (around 140inches), 370 cm (around 145 inches), 380 (around 150 inches) cm or 390cm (around 155 inches). The fan diameter may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds).

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 250 cm to 300 cm (for example 250 cm to 280 cm) may be inthe range of from 1700 rpm to 2500 rpm, for example in the range of from1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100rpm. Purely by way of further non-limitative example, the rotationalspeed of the fan at cruise conditions for an engine having a fandiameter in the range of from 320 cm to 380 cm may be in the range offrom 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to1800 rpm, for example in the range of from 1400 rpm to 1600 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.3, 0.31,0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in thisparagraph being Jkg⁻¹K⁻¹/(ms⁻¹)²). The fan tip loading may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds).

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, or 17. The bypass ratio may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The bypass duct may besubstantially annular. The bypass duct may be radially outside the coreengine. The radially outer surface of the bypass duct may be defined bya nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds).

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹ s, 105 Nkg⁻¹ s, 100 Nkg⁻¹ s, 95 Nkg⁻¹ s, 90 Nkg⁻¹ s, 85 Nkg⁻¹ s or80 Nkg⁻¹ s. The specific thrust may be in an inclusive range bounded byany two of the values in the previous sentence (i.e. the values may formupper or lower bounds). Such engines may be particularly efficient incomparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). The thrust referred to abovemay be the maximum net thrust at standard atmospheric conditions at sealevel plus 15 deg C. (ambient pressure 101.3 kPa, temperature 30 degC.), with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). The maximum TET may occur, for example, at a high thrustcondition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a blisk or a bling. Any suitable method may be used tomanufacture such a blisk or bling. For example, at least a part of thefan blades may be machined from a block and/or at least part of the fanblades may be attached to the hub/disc by welding, such as linearfriction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 16, 18, 20, or 22 fan blades.

As used herein, cruise conditions may mean cruise conditions of anaircraft to which the gas turbine engine is attached. Such cruiseconditions may be conventionally defined as the conditions atmid-cruise, for example the conditions experienced by the aircraftand/or engine at the midpoint (in terms of time and/or distance) betweentop of climb and start of decent.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be the cruise condition. For someaircraft, the cruise conditions may be outside these ranges, for examplebelow Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions at an altitude that is in the range offrom 10000 m to 15000 m, for example in the range of from 10000 m to12000 m, for example in the range of from 10400 m to 11600 m (around38000 ft), for example in the range of from 10500 m to 11500 m, forexample in the range of from 10600 m to 11400 m, for example in therange of from 10700 m (around 35000 ft) to 11300 m, for example in therange of from 10800 m to 11200 m, for example in the range of from 10900m to 11100 m, for example on the order of 11000 m. The cruise conditionsmay correspond to standard atmospheric conditions at any given altitudein these ranges.

Purely by way of example, the cruise conditions may correspond to: aforward Mach number of 0.8; a pressure of 23000 Pa; and a temperature of−55 deg C.

As used anywhere herein, “cruise” or “cruise conditions” may mean theaerodynamic design point. Such an aerodynamic design point (or ADP) maycorrespond to the conditions (comprising, for example, one or more ofthe Mach Number, environmental conditions and thrust requirement) forwhich the fan is designed to operate. This may mean, for example, theconditions at which the fan (or gas turbine engine) is designed to haveoptimum efficiency.

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

DESCRIPTION OF THE DRAWINGS

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4 shows schematically a perspective view from the rear of an enginewithout its nacelle and without most of its inner cowl;

FIG. 5 shows schematically at left a transverse cross-section throughthe engine of FIG. 4 when door sections of its inner cowl are closed,and at right a transverse cross-section through the engine when thesedoor sections are swung open;

FIG. 6 shows a side perspective view of front end portions of a keelbeam and lower splitter of the engine of FIG. 4;

FIG. 7 shows a top perspective view of the front end portions of thekeel beam and lower splitter of the engine of FIG. 4;

FIG. 8 shows a cooling flow on a schematic side view of the enginewithout its nacelle and without most of the inner cowl; and

FIG. 9 shows the cooling flow on a perspective view looking down on tothe keel beam of the engine.

DETAILED DESCRIPTION

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine is mounted to an airframe, e.g. under a wing, by amounting pylon 46. The engine 10 comprises an air intake 12 and apropulsive fan 23 that generates two airflows: a core airflow A and abypass airflow B. The gas turbine engine 10 comprises a core 11 thatreceives the core airflow A. The engine core 11 comprises, in axial flowseries, a low pressure compressor 14, a high-pressure compressor 15,combustion equipment 16, a high-pressure turbine 17, a low pressureturbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gasturbine engine 10 and defines a bypass duct 22 and a bypass exhaustnozzle 18. The pylon 46 forms a bifurcation in the bypass duct where ittraverses the duct to join to the engine core 11. The bypass airflow Bflows through the bypass duct 22, where it is straightened by a row ofouter guide vanes 40 before exiting the bypass exhaust nozzle 18.Rearward of the outer guide vanes 40, the engine core 10 is surroundedby an inner cowl 41 which provides an aerodynamic fairing defining aninner surface of the bypass duct 22. The fan 23 is attached to anddriven by the low pressure turbine 19 via a shaft 26 and an epicyclicgearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 constrains the planet gears32 to precess around the sun gear 28 in synchronicity whilst enablingeach planet gear 32 to rotate about its own axis. The planet carrier 34is coupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Radially outwardly of the planet gears 32 andintermeshing therewith is an annulus or ring gear 38 that is coupled,via linkages 40, to a stationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2. For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 20, 22meaning that the flow through the bypass duct 22 has its own nozzle thatis separate to and radially outside the core engine nozzle 20. However,this is not limiting, and any aspect of the present disclosure may alsoapply to engines in which the flow through the bypass duct 22 and theflow through the core 11 are mixed, or combined, before (or upstream of)a single nozzle, which may be referred to as a mixed flow nozzle. One orboth nozzles (whether mixed or split flow) may have a fixed or variablearea. In some arrangements, the gas turbine engine 10 may not comprise agearbox 30.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

FIG. 4 shows schematically a perspective view from the rear of theengine 10 with the nacelle 21, most of the inner cowl 41 and the pylon46 removed. A fan case 42 defines an outer surface of the bypass duct 22and towards the rear of the fan case an inner ring 47 defines an innersurface of the bypass duct 22. The outer guide vanes 40 extend radiallyfrom the inner ring to the fan case, and the engine core 11 projectsrearwardly from the plane of the outer guide vanes. A fixture 45 locatedat top dead centre behind the inner ring provides a connection point forthe mounting pylon 46 which mounts the engine to the airframe.

The inner cowl 41 includes two door sections, one on either side of theengine 10, with each door section being pivotably openable about arespective pivot line which extends from front to back along that doorsection's side of the mounting pylon 46. This allows the door sectionsto be swung upwards and away from the engine core 11 for maintenanceaccess. Conveniently, the top edges of the door sections can form thepivot lines. FIG. 5 shows schematically at left a transversecross-section through the engine when the door sections of the innercowl 41 are closed (and also door sections of the nacelle 21 surroundingthe fan case 42 are closed), and at right a transverse cross-sectionthrough the engine when these door sections are swung open.

Returning to FIG. 4, a two-part inner barrel 43 attached to the innerring 47 bridges the space between the inner ring and the inner cowl, andprovides circumferentially extending V-grooves at its rear edge forengaging the front edges of the door sections when they are closed. Inaddition, a keel beam 48 extends rearwardly from bottom dead centre ofthe inner ring 47 to provide latching formations for latching to loweredges of the door sections when they are closed. The keel beam 48 canhave one or more cut-out drain holes (not shown) to provide core zoneliquid drainage, and a profiled drainage guidance surface to provide apositive drainage force via gravity to the drain holes.

A lower splitter 49 may also traverse the bypass duct, and conveniently,the keel beam can then form an interface for mounting a radially innerend of the lower splitter. This mounting arrangement is advantageouswhen the engine has a large volume of core-mounted accessories in thespace between the engine core and the keel beam such that a direct mountof the lower splitter to a casing of the engine core is impractical. Thelower splitter can be fastened to the keel beam by a straightforwardbolting arrangement positioned where the components are in contact.

FIGS. 6 and 7 show side and top perspective views of front end portionsof the keel beam 48 and lower splitter 49. They also show engineaccessories 40 (drawn semi-transparently in FIGS. 6 and 7) which arelocated above the keel beam within an engine zone bounded on a radiallyouter side by the inner cowl 41, and on a radially inner side by corecasings 44 of the engine core 11. The engine accessories 40 include anaccessory gearbox driven by a take-off (such as a radial drive shaft—notshown) from the core shaft 26. Other accessories may include any one ormore of a power generator, a fuel pump, an oil pump, a hydraulic pump,and an engine starter motor. The engine accessory gearbox has a frontend that receives the drive from the drive shaft 26 and has a train ofspur gears which transfer the drive to the other accessories. These spurgears are arranged in a line along a central spine member 52, with theother accessories projecting from opposite sides of the spine member.The central spine member extends substantially parallel to the engineaxis 9 with the rotation axes of the spur gears perpendicular to theengine axis.

The engine zone in which the engine accessories 40 are located can reachair temperatures above 300° C. Thus a ventilation system is provided tomaintain engine accessories within acceptable thermal limits. Thissystem is located on the keel beam 48 and comprises a ventilation inlet60 which faces forwards at engine bottom dead centre in the bypass duct22 to receive a portion of the relatively low temperature (<100° C.)bypass airflow B as a cooling flow for the accessories. Kinetic energyof the received bypass flow is converted at the inlet into an increasedpressure of the cooling flow, whereby the cooling flow has a pressuredifferential relative to air pressure in the engine zone which issufficient to keep the accessories ventilated even at low engine powerconditions. To direct the cooling flow where it is needed, theventilation system further comprises a manifold 61 located on theradially inner surface of the keel beam in fluid communication with theventilation inlet to collect the cooling flow therefrom. The coolingflow leaves the manifold through plural exhaust holes 62 to impinge uponand extract heat from the engine accessories, protecting the accessoriesfrom the heat present in the engine zone. For example, the manifold andits exhaust holes can be configured as shown so that the holes form twolines which extend substantially parallel to the engine axis 9 toopposite sides of the accessories. In this way, the cooling air can beexhausted at high velocities over the full length of the engineaccessories 40.

Optionally, the ventilation system also comprises a control unit 63which is under the control of the engine's EEC, and which regulates theamount of cooling flow required to ventilate the engine accessories 40at different flight conditions thereby modulating and optimising thecooling flow at different flight conditions, to reduce its impact onengine performance and specific fuel consumption. Conveniently, thecontrol unit can be located between the inlet 60 and the manifold 61 tomediate transfer of the cooling flow therebetween. However, particularlyif regulation of the cooling flow has a little impact on engineperformance, the control unit can be omitted to reduce weight andcomplexity of the ventilation system.

FIG. 8 shows a schematic side view of the engine with the nacelle 21,most of the inner cowl 41 and the pylon 46 removed, and FIG. 9 shows aperspective view looking down on to the keel beam 48. The bypass airflowB received into the inlet 60, and the cooling flow through the manifold61 and out of the exhaust holes 62 are indicated by arrowed lines andillustrate how the cooling flow surrounds the engine accessories 40 in acontinuous film of cooled air.

The ventilation inlet 60 takes advantage of the dynamic head of thebypass airflow B to generate a sufficient driving force for the flow totravel through the control unit 63 and the manifold 61. For example, theventilation inlet can be a total pressure inlet, e.g. converting atleast 50% of the kinetic energy of the received bypass airflow into apressure rise, and preferably substantially stagnating the receivedbypass airflow.

Although shown in FIGS. 6 to 9 with just one ventilation inlet 60, theventilation system may have plural inlets. Additionally, oralternatively, the inlet or inlets can be located elsewhere, such as onthe lower splitter 49.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

1. A gas turbine engine for an aircraft, the engine including: an enginecore comprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor; core casings surrounding the engine core; anaerodynamic owl which surrounds the core casings; a propulsive fanlocated upstream of the engine core, the fan generating a core airflowwhich enters the core engine and a bypass airflow which enters a bypassduct surrounding the aerodynamic cowl; and one or more engineaccessories mounted in a space between the core casings and theaerodynamic cowl; wherein the gas turbine engine further includes: oneor more ventilation inlets which receive a portion of the bypass airflowas a cooling flow for the engine accessories, the ventilation inletsbeing configured to convert at least a portion of the kinetic energy ofthe bypass flow received therein into a pressure rise such that thepressure of the cooling flow is increased relative to the pressure ofthe bypass airflow; and a manifold in fluid communication with theventilation inlets to collect the cooling flow therefrom, the manifoldhaving plural exhaust holes therefrom through which the cooling flowleaves the manifold to impinge upon and thereby cool the engineaccessories.
 2. A gas turbine engine according to claim 1, wherein theone or more ventilation inlets convert at least 50% of the kineticenergy of the portion of the bypass airflow received therein into apressure rise.
 3. A gas turbine engine according to claim 1, whichfurther includes a control unit which is operable to regulate the amountof cooling flow collected by the manifold from the ventilation inlets.4. A gas turbine engine according to claim 1, wherein the exhaust holesare arranged such that the cooling flow impinging upon the engineaccessories forms a continuous film of cooling air around theaccessories.
 5. A gas turbine engine according to claim 1, wherein theexhaust holes are formed into two lines which extend substantiallyparallel to the principal rotation axis of the engine to opposite sidesof the engine accessories.
 6. A gas turbine engine according to claim 1,wherein the one or more engine accessories are mounted verticallybeneath the core casings.
 7. A gas turbine engine according to claim 1,wherein the one or more engine accessories include an engine accessorygearbox driven by a take-off from the core shaft.
 8. A gas turbineengine according to claim 7, wherein the engine accessory gearboxincludes a train of spur gears which transfer the drive to other engineaccessories, the spur gears being arranged in a line and having axes ofrotation which extend perpendicularly to the principal rotation axis ofthe engine.
 9. A gas turbine engine according to claim 8, wherein thetrain of spur gears is mounted along a central spine member, the otherengine accessories projecting from opposite sides of the spine member.10. A gas turbine engine according to claim 1, which further includes acircumferential row of outer guide vanes located in the bypass ductrearwards of the propulsive fan, the outer guide vanes extendingradially outwardly from an inner ring which defines a radially innersurface of the bypass duct; wherein the aerodynamic cowl is rearwards ofthe inner ring and includes two door sections located on respective andopposite sides of the engine, each door section being pivotably openableabout an upper edge thereof to enable maintenance access to the enginecore, the aerodynamic cowl further including a keel beam which extendsrearwardly from the inner ring at bottom dead centre thereof to providelatching formations for latching to a lower edge of each door sectionwhen it is closed; and wherein the ventilation inlets are located on theradially outer surface of the keel beam.
 11. A gas turbine engineaccording to claim 10, wherein the manifold is located on the radiallyinner surface of the keel beam.
 12. A gas turbine engine according toclaim 1, further including a power gearbox that receives an input fromthe core shaft and outputs drive to the propulsive fan so as to drivethe fan at a lower rotational speed than the core shaft.
 13. A gasturbine engine according to claim 1, wherein the turbine is a firstturbine, the compressor is a first compressor, and the core shaft is afirst core shaft; the engine core further includes a second turbine, asecond compressor, and a second core shaft connecting the second turbineto the second compressor; and wherein the second turbine, secondcompressor, and second core shaft are arranged to rotate at a higherrotational speed than the first core shaft.